Method to restore an airfoil leading edge

ABSTRACT

The present invention provides methods and apparatus to restore a blade leading edge on a gas turbine engine component such as an airfoil of a turbine blisk. The method utilizes welding image technology and power control systems in order to provide effective welding with superalloy materials such as Inconel 713 and Inconel 625. The method includes machining away a damaged leading edge and providing a repaired region through successive depositions of superalloy powder filler through laser fusion welding. Deposition material is added until the repaired region exceeds the original dimensions of the airfoil. The airfoil is then machined and finished to return it to original airfoil dimensions.

FIELD OF THE INVENTION

The present invention relates to laser welding. Additionally theinvention relates to the apparatus and techniques used to repair theleading edge of airfoils that have suffered degradation or wear. Moreparticularly, the invention relates to a method to restore, by laserwelding techniques with powder filler, the leading edge on the blades ofturbine blisks that have been eroded by foreign object damage.

BACKGROUND OF THE INVENTION

Turbine engines are used as the primary power source for many types ofaircraft. The engines are also auxiliary power sources that drive aircompressors, hydraulic pumps, and industrial gas turbine (IGT) powergeneration equipment. Further, the power from turbine engines is usedfor stationary power supplies such as backup electrical generators forhospitals and the like.

Most turbine engines generally follow the same basic power generationprocedure. Compressed air generated by axial and/or radial compressorsis mixed with fuel and burned, and the expanding hot combustion gasesare directed against stationary turbine vanes in the engine. The vanesdeflect the high velocity gas flow so as to impinge on the turbineblades mounted on a rotatable turbine disk. The force of the impinginggas causes the turbine disk to spin at high speed. Jet propulsionengines use the power created by the rotating turbine disk to draw moreair into the engine, and the high velocity combustion gas is passed outof the gas turbine aft end to create forward thrust. Other engines usethis power to turn one or more propellers, fans, electrical generators,or other devices.

In an attempt to increase the efficiencies and performance ofcontemporary gas turbine engines generally, engineers have progressivelypushed the engine environment to more extreme operating conditions. Theharsh operating conditions of high temperature and pressure that are nowfrequently specified place increased demands on enginecomponent-manufacturing technologies and new materials. Indeed thegradual improvement in engine design has come about in part due to theincreased strength and durability of new materials that can withstandthe operating conditions present in the modern gas turbine engines. Withthese changes in engine materials, there has arisen a corresponding needto develop new repair methods appropriate for such materials.

One development in the design of gas turbine engine components has alsobeen the introduction of integrally structured airfoil and rotordevices, blisks. A turbine blisk, for example, includes turbine airfoilsthat are integrally formed with the perimeter of a rotor disk by, forexample, integral casting. This design provides the advantage ofeliminating the connection between individual airfoils and the rotor ata dovetail. The blisk, by having a unitary construction, also provides astrong mechanical connection between the airfoil region and the rotordisk region thereby allowing for a more efficient positioning of theairfoils. This results in an improved performance of the blisk in termsof weight and component size.

The development of the blisk as a gas turbine engine component haspresented challenges with respect to repair strategies. Individualairfoils are now permanently attached to the rotor disk, which meansthat damaged airfoils cannot easily be removed for repair, as has beendone with individual turbine blades. Nonetheless, blisks do have anormal life cycle and must be repaired or replaced at the end. Blisksare impacted by foreign objects such as sand, dirt, and other suchdebris. Blade leading edge damage, for example, is a common failureexperienced in blisks. The leading edge is subject to foreign objectdamage or erosion after a period of service time.

The option of throwing out worn engine components such as turbine blisksand replacing them with new ones is not an attractive alternative.Blisks are very expensive due to costly material and manufacturingprocesses. Consequently there is a strong financial need to find anacceptable and efficient repair method for turbine blisks.

Turbine blisks used in modern gas turbine engines are frequentlycastings from a class of materials known as superalloys. The superalloysinclude nickel-based, cobalt-based and iron-based superalloys. Inconel713 is a typical superalloy used in blisk construction. In the castform, turbine blisks made from advanced superalloys include manydesirable properties such as high elevated-temperature strength and goodenvironment resistance. Advantageously, the strength displayed by thismaterial remains present even under stressful conditions, such as hightemperature and high pressure, experienced during engine operation.

Disadvantageously, the superalloys generally are very difficult to weldsuccessfully. Traditional repair methods have proven less thansatisfactory for superalloy materials. For example, with regard toturbine blades, as opposed to turbine blisks, known welding techniquesoften include heating a turbine blade to high temperatures, ranging from1800° F. to 2000° F. before the welding process. However, at such anelevated temperature the turbine blade may experience heat cracking andfracturing, rendering the blade unusable for further engine service.Other welding techniques similarly suffer from a lack of thermal controland provide too much localized heat during welding to render aneffective repair with superalloy blisk airfoils. Superalloys aresusceptible to microcracking during localized heating. Moreover, thecomplex geometry of the airfoil, and particularly, the shape of theleading edge, makes it difficult to deposit filler or cladding materialthereon. Finally, the turbine blisk airfoil has a region thatexperiences high stress. It has proven difficult to provide filler orcladding material across a high stress region with sufficient strengthand adherence such that the airfoil can be returned to service. Thusprevious repair strategies used on blisks have avoided the high stressregion.

Hence, there is a need for a turbine airfoil restoration method thataddresses one or more of the above-noted drawbacks and needs. Namely, arepair method is needed that can fully restore geometry, dimension anddesired properties of degraded turbine blisk airfoils and/or a methodthat allows control of welding parameters so that blisk repairs may beaffected without heat cracking and damage to the airfoil and/or a methodthat allows for repairs across the high stress zone of a turbine blisk.Finally, it would be desired to provide a method to restore a bliskairfoil that, by virtue of the foregoing, is therefore less costly ascompared to the alternative of replacing worn parts with new ones. Thepresent invention addresses one or more of these needs.

SUMMARY OF THE INVENTION

The present invention provides an apparatus and methods for use inrestoring turbine blisk airfoils through laser welding techniques. Inone embodiment, the invention provides a powder-fed CO₂ laser welderthat is capable of welding superalloy filler material to the superalloysubstrate of the blisk. The damaged leading edge of the turbine blisk iscut back, and filler material is welded into the leading edge area. Thesurface contour of the blisk is then restored to a desired geometry.

In one embodiment, and by way of example only, there is provided amethod for resurfacing the leading edge of an airfoil comprising thesteps of: removing material from the leading edge of an airfoil;preparing the airfoil for welding; selecting a weld path using an imagesystem; determining welding parameters in order to avoid cracking; andlaser cladding filler material onto the airfoil. The laser cladding offiller material may take place in a high stress region of the airfoil.

In a further embodiment, still by way of example only, there is provideda method for resurfacing the damaged leading edge of a turbine bliskairfoil comprising the steps of: machining material away from a damagedleading edge to a selected height and depth; inspecting the machinedarea by fluorescent penetrating inspection; preparing the leading edgefor welding; determining a weld path with a laser image system; andperforming a laser fusion welding with a superalloy powder filler and aCO₂ laser. The laser fusion may use a coaxial powder feeder nozzle. Thelaser fusion welding may take place across a high stress region of anairfoil. The method may include measuring the depth of the depositionand repeating a laser fusion welding until a desired thickness isachieved. The superalloy powder filler may comprise Inconel 625superalloy powder, and the substrate may be composed of Inconel 713. Themachining step may include machining material to a selected height anddepth so as to remove damaged portions of the leading edge. Inspectingthe airfoil may further comprise inspecting in order to confirm theabsence of cracks that would disqualify the airfoil from repair. Themethod may also include machining a repaired airfoil to a desiredcontour.

In still a further embodiment, and still by way of example only, thereis provided a resurfaced airfoil comprising: an airfoil with a leadingedge, a trailing edge, and a top edge integrally connected to a blisk; asubstrate region of the airfoil; a repaired region of the airfoil laserwelded by powder fusion repair to the substrate region wherein therepaired region extends from a welding surface to a leading edge of theairfoil and from a welding surface to a top edge of the airfoil. Therepaired region may be formed by overlapping laser cladding depositionsof powdered alloy, and the repaired region may cross a high stressregion of the airfoil. Material of the repaired region may comprise asuperalloy such as Inconel 625. Material of the substrate region maycomprise a superalloy such as Inconel 713. The substrate region mayfurther comprise a weld surface, which may be arcuate shaped, at whichthe repaired region is welded to the substrate region.

Other independent features and advantages of the method to restore anairfoil leading edge will become apparent from the following detaileddescription, taken in conjunction with the accompanying drawings whichillustrate, by way of example, the principles of the invention.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 is a perspective view of a turbine blisk airfoil that may berestored according to an embodiment of the present invention.

FIG. 2 is a perspective view of a laser welding system that may be usedto perform airfoil restoration according to an embodiment of the presentinvention.

FIG. 3 is a perspective view of a turbine blisk airfoil with damagedarea machined away according to an embodiment of the present invention.

FIG. 4 is a flow chart that illustrates steps in an exemplary embodimentof the method to restore an airfoil leading edge.

FIG. 5 is a photomicrograph showing a substrate region and a repairedregion of a restored airfoil according to an embodiment of the presentinvention.

FIG. 6 is a photomicrograph showing a substrate region and a repairedregion of a restored airfoil according to an embodiment of the presentinvention.

FIG. 7 is a photomicrograph showing a substrate region and a repairedregion of a restored airfoil according to an embodiment of the presentinvention.

FIG. 8 is a photomicrograph showing a substrate region and a repairedregion of a restored airfoil according to an embodiment of the presentinvention.

DETAILED DESCRIPTION OF A PREFERRED EMBODIMENT

The following detailed description of the invention is merely exemplaryin nature and is not intended to limit the invention or the applicationand uses of the invention. Furthermore, there is no intention to bebound by any theory presented in the preceding background of theinvention or the following detailed description of the invention.Reference will now be made in detail to exemplary embodiments of theinvention, examples of which are illustrated in the accompanyingdrawings. Wherever possible, the same reference numbers will be usedthroughout the drawings to refer to the same or like parts.

A typical airfoil 10 of a turbine blisk is illustrated in FIG. 1. Such ablade may have a different geometric and dimension design, depending onthe engine model and its application. For a typical aero-engine, aturbine blisk airfoil is typically a few inches in length. Airfoil 10 ischaracterized by a complex geometry that changes in three dimensions. Agas turbine airfoil may be welded to, or cast in unitary form, with hub11 shown in partial view. In an engine assembly multiple such turbineairfoils are positioned in adjacent circumferential position along thehub or rotor disk. Multiple blisks or other rotor assemblies may besequentially positioned in the engine. Airfoil 10 is a cuplike structurethat includes a concave face 13 and a convex face (not shown) on thereverse side of the airfoil. Airfoil 10 extends radially outwardly fromthe hub. A top edge 12 defines the radial end of the airfoil.

In operation, gases impinge on concave face 13 of airfoil 10 therebyproviding the driving force for the turbine engine. Pressure develops onconcave face 13 while suction develops on the convex face. This forceacting on the airfoil thereby spins hub 11. Turbine airfoil 10 alsoincludes leading edge 17 and trailing edge 18 which represent the edgesof the airfoil that firstly and lastly encounter an air stream passingaround it. Leading edge 17 is subject to wear and degradation. Partlythis arises from debris and contaminants carried in the airstream. Thisdebris impacts leading edge at high velocity thus leading to nicks,wear, and erosion, all of which impair the engine performance.

Referring now to FIG. 2 there is shown a schematic diagram of a generalapparatus for laser generation and control that may be used in airfoilrestoration according to an embodiment of this invention. Lasergenerating means 20 generates a laser used in the welding system. Alaser is directed through a laser conveyance which may include passingthe laser through beam guide 21, through mirror 22, and through focuslens 23. In some embodiments, beam guide 21, mirror 22, and focus lens23 may not be present, or may have different configurations. The laserthen impinges on a surface of a work piece 24. Components such as beamguide 21, mirror 22, and focus lens 23 are items known in the art oflaser powder fusion welding. Beam guide 21 may include fiber opticmaterials such as fiber optic laser transmission lines.

A means for providing a filler or cladding material is also included foruse with the laser. Preferably this filler material may be provided inpowder feeder 25. In such an embodiment the powder is fed onto theworkpiece 24 through powder feed nozzle 26. A coaxial or off-axisarrangement may be used with powder feed nozzle 26 with respect to themain laser. Alternatively, filler material may be provided through othermeans such as a wire feed.

Other components of the laser welding system include a vision CCD camera27 and video monitor 28. The image taken by the camera 27 can also befedback to the controller screen 30 for positioning and weldingprogramming. Controller 30 is similarly connected to operable pieces ofthe welding system and thereby controls features such as welding power,energy power (on/off and pulse/continuous) laser beam size, weld path,welding velocity, and filler delivery. The workpiece 24 is held on awork table 29. An inert gas shield (not shown) is fed through guides(not shown) onto the workpiece 24. The inert gas shield is directed ontoa portion of the surface of the workpiece 24 during laser welding.

Controller 30 preferably includes a computer numerically controlled(CNC) positioning and digital imaging system. CNC controller 30coordinates components of the system and allows for automated,programmed welding. The controller 30 also guides movement of the laserand powder feed across the face of the workpiece 24. The controller 30thus allows for a fully automated laser welding operation. Moreover, theimaging and vision aspects of the controller allow it to selectweldpaths (and welding parameters) so as to minimize or effectivelyeliminate stress and heat-related damage to the workpiece. In apreferred embodiment, movement of the workpiece in the XY plane isachieved through movement of the worktable. Movement in the up and down,or Z-direction is achieved by control of the laser arm; i.e., pulling itup or lowering it. Alternative methods of control are possible, such ascontrolled movement of the workpiece in all three directions, X, Y, andZ as well as rotation and tilt.

It will be appreciated that the geometric configuration of airfoils on ablisk can lead to space limitations with respect to welding machinery.The components of the welding system that come into close contact withthe blade leading edge must be sufficiently compact so that physicalmaneuvering is possible. Thus, in a preferred embodiment, a CO₂ laserwith a concentric powder nozzle is selected as the close contact lasersystem. This permits the application of sufficient material and energyin the direction of a blade leading edge to affect repairs.

A damaged airfoil such as illustrated in FIG. 1 can be restored to adesired shape using, for example, an apparatus as illustrated in FIG. 2.In a first step of the repair process, damaged material on an enginecomponent is machined. For example damaged leading edges of a blisk aremachined so as to remove the damaged portion. The remaining airfoilmaterial should not suffer from any degree of damage that would preventa restored airfoil from returning to service. In a preferred embodiment,the leading edges 17 (or other damaged area such as a tip or top edge)are machined back to predetermined limits. The limits may refer to adegree of machining in a lateral direction starting from the leadingedge 17 and a depth direction starting the top edge 12 of the blade. Thepredetermined limits provide a margin of safety whereby any damagedmaterial is removed. When an automated machining operation is to beused, it is preferred that all airfoils be machined to the same limits.The limits of machining may be determined by an inspection step of theairfoils. Known methods of machining or grinding may be used for thematerial removal. It is preferably done by automated means using amulti-axis numerically controlled milling machine. In one embodiment,digital information regarding the blade's blueprint or actual geometryis used to program the desired machining operation.

Referring now to FIG. 3 there is shown an airfoil after a machining. Theportion of the airfoil that has been cut away leaves a generally arcuateshaped area on the substrate region 31 of the blade, which is theremaining mass of the airfoil. The cut away portion also reveals a weldsurface 32 on substrate region. Weld surface 32 extends from leadingedge 17 to top edge 12.

The machined airfoil may optionally receive an inspection, such as afluorescent penetration inspection. This inspection can determinewhether the substrate region 31 has any imperfections that woulddisqualify the blade from service even after repair. Additionally, theinspection can confirm that all damaged material has been removed. Oncematerial has been removed, the area of the airfoil that is now exposedmay also be prepared for welding. This may include standard treatmentssuch as grit blasting and solution treatment.

In a further step filler material is deposited by laser weldingtechniques on welding surface 32. Preferably, filler material isdeposited through the use of powder fusion welding. In this system,filler material in powder form is discharged so that it is melted by thelaser beam and welds on the desired surface of the workpiece. Weld pathsare chosen to avoid stress-concentration areas. During welding a singleweld bead is preferred. However, if the bead is not of sufficientdimension to cover the entire machined surface, then a stitch patternmay be used to provide a desired weld build-up as well as area coverage.Processing parameters are also chosen to control thermal input duringthe welding operation. It is preferred to minimize the amount of heatdischarged through the laser to the minimum amount necessary to affectlaser welding. Further, the area of the laser beam spot and laservelocity are similarly determined in order to regulate the heatexperienced by the substrate and the stress caused by the laser weldingprocess. Thus, the welding operation avoids microcracking in the weldarea and the heat affected zone.

In a preferred embodiment, the power of the laser projected onto thewelding area 32 is between about 50 to about 2500 watts and morepreferably between about 50 to about 1500 watts. The powder feed rate ofpowder filler material is between about 1.5 to about 20 grams per minuteand more preferably about 1.5 to about 10 grams per minute. Travelingspeed for the motion of the substrate work table 29 relative to thelaser beam is about 3 to about 22 inches per minute and more preferablyabout 5 to about 14 inches per minute. The size of the laser beam spotprojected onto the welding area 32 is about 0.02 to about 0.1 inches indiameter and more preferably about 0.04 to about 0.06 inches. Thelaser-welded bead width that results through the laser welding is thusabout 0.02 to about 0.100 inches and more preferably about 0.04 to about0.06 inches in width.

Multiple passes may be used to build up a required dimension of materialwhere one pass overlaps a previous pass and successive passes are laidatop a previous pass. Similarly, the method allows for cladding of anarea greater than that covered in a single pass by laying successivepasses alongside previous passes thus covering a desired area. Ifneeded, repetitions of the laser welding passes can be done in order toachieve a required level of buildup and/or coverage over a requiredarea; this is accomplished by depositing successive layers of fillermaterial on top of one another. Upon conclusion of a first pass, orother passes, the controller can check the thickness of the welddeposit. If needed, additional weld deposits can then be conducted.

Laser welding depositions continue until sufficient material has beendeposited. Sufficient material is deposited when the deposited materialnow occupies the volume of material that had been machined away from theairfoil. Thus, deposited material preferably extends to a point beyondleading edge 17 and top edge 12 of the airfoil in its originalcondition. The result is a mass of newly deposited material thatoccupies a repaired region. The material of the repaired region isfusion welded with the material of the substrate region. Further theweld is characterized by a lack of any degree of cracks, voids, ordiscontinuities that would disqualify the turbine blisk from service.

The powder or filler used in the laser welding process is compatiblewith the alloy comprising the workpiece. The dimension of filler powderis measured by its mesh size. Preferred powder size ranges from +100mesh to −325 mesh. In one preferred embodiment, Inconel 625 powder isused as a filler material to restore an airfoil whose substrate is madeof Inconel 713. Some superalloy filler materials that are also suitablefor the practice of this invention and that are commercially availablein powder and wire form include: HS188, Stellite 694, Hastelloy X,Inconel 625, INCO 738, INCO 939, MarM247, Rene 80, and C 101. Somematrix or base superalloys, which are suitable for the practice of thisinvention and may be laser welded include: Inconel 713, INCO738, C101,MarM-247, Rene80, GTD111, Rene125, Rene142, SC 180, Rene N5 and N6,CMSX-2, CMSX-4 and CMSX-10, and PWA 1480 and 1484.

INCONEL is a trade name owned by Inco Alloys International, Inc. Thename INCONEL refers to a number of nickel-based superalloys. Several ofthe Inconel superalloys are used in aerojet applications. The same orsimilar superalloys may be manufactured by sources which may use adifferent name. Inconel 625 and 713, which are preferred alloys for usein the present invention, have the following general compositions:INCONEL 625 INCONEL 713 Wt % Wt. % Element Composition ElementComposition Carbon  0.1 max Carbon  0.20 max. Manganese  0.5 maxManganese  1.0 max. Sulphur 0.015 max Sulphur 0.015 max. Silicon  0.5max. Silicon  1.0 max. Chromium   20-23 Chromium 11.00-14.00 Molybdenum  8-10 Molybdenum  3.5-4.5 Titanium  0.4 max Titanium  0.25-1.25Aluminum  0.4 max Aluminum  5.5-6.5 Iron    5 max. Iron  5.0 max. Cobalt   1 max. Co + Ta  1.0-3.0 Niobium 3.15-4.15 Nickel remainderPhosphorous  0.15 max Nickel remainder

After completion of the laser powder fusion step, the airfoil may bemachined and finished so as to return the airfoil to a desired shape orgeometry. A rough machining might be necessary to remove abundant weldmetal prior to a final machining. A final machining may then beperformed by hand blending or a CNC milling operation. A preferredgeometry is the blueprint geometry of the original airfoil, although, asis understood in the art, approximations to this shape are acceptable.In a preferred embodiment, material in the repaired region is initiallyoverdeposited with respect to the starting boundaries of the airfoil,the leading edge and the top edge. However, it is also preferred not tounnecessarily overdeposit material as this leads to wastage of materialand further processing to restore the airfoil to a desired shape.

Referring now to FIG. 4 there is shown a flowchart that illustratessteps in a preferred embodiment of the method to restore an airfoilleading edge. In step 41 an airfoil is machined to remove damagedmaterial. In step 42 the machined airfoil optionally receives aninspection to confirm the removal of the damaged portion of the airfoil.In step 43, filler material is deposited onto the airfoil. And in step44 the airfoil is machined to a desired shape.

The following example is illustrative of the principles of theinvention. A set of turbine rotor blades were selected for leading edgerestoration. The selected blades had been in service in a third stageturbine rotor assembly of the GTCP331-250, a Garret Auxiliary Power Unit(APU). The blades had been subjected to FOD (foreign object damage)through routine usage. The base metal of the turbine blades was Inconel713. The weld filler material was Inconel 625.

In a first step the blades were cleaned by soaking them in an alkalinesolution and subjecting the blades to a vapor blasting.

The blade leading edge was machined. The machining was to a depth andheight so as to removes nicks, wear, and erosion damage on the bladeleading edge. The machining extended to a maximum of 0.500″ in lengthdown the leading edge from the existing tip height and to a maximum of0.080″ in depth back from the leading edge. The length and depth ofmachining was chosen such that all regions of wear, damage, and erosionwere removed.

The blades received a visual and Fluorescent Penetrating Inspection(FPI) to determine reparability. Cracks or other imperfections thatwould disqualify the blade from service were noted. This step confirmedthat all surface defects were removed after machining.

The surface areas to be welded received a welding preparation. Thisincluded grit blasting the surface areas with 220 grit aluminum oxide at40 psi.

The blades then received a laser cladding on the machined leading edge.The laser welding used an Inconel 625 alloy powder as the fillermaterial. A Huffman 205 CO₂ laser system was used to perform the lasercladding operation.

The leading edge of the blades then received both a rough machining anda finish blend machining to remove excessive weld material. The finishblending also restored the blade contour to an original, orapproximately original, profile.

The blades were also subjected to a heat treatment with stress reliefcycle.

Next the welded area received a chemical etch. An etchant such as ferricchloride solution may be used. The etching allows for an accurate FPIreading. The blades then received an FPI inspection at the weldedleading edge. The FPI inspection was in accordance with ASTM-E1417, TypeI, Method D, and Sensitivity level 4. No cracks were identified.

At this point in the procedure, the blades are restored. They may bereturned to service. However, as a further examination, the restoredblades were cut up and subjected to a metallurgical analysis in order toevaluate the quality of the repair.

The blades were sectioned according to a plan that would allow forevaluation of the weld along multiple axes. Cut up plan A followed atransverse cross-section at different heights, A₁ and A₂. Cut up plan Bfollowed the edge of the weld, a vertical cut along the blade. Cut upplan C followed a cord length of the airfoil. The mounts that resultedfrom the sectioning were polished with 0.05 micron silica and etchedwith Kalling's reagent. The mounts were polished three times, except forthe Section C mounts, which did not have adequate metal for threepolishes. The results of the metallurgical inspection are shown in thefollowing table. Photomicrographs of the mounts are shown in FIGS. 5through 8. Material Metallurgical Cut-Up Plan Polishes Removal Results C1 0.019″ OK 2 0.012″ OK C 1 0.022″ OK (See FIG. 5) 2 0.010″ OK A 10.015″ OK 2 0.021″ OK (See FIG. 6) 3 0.034″ OK A 1 0.017″ OK (See FIG.7) 2 0.009″ OK 3 0.024″ OK A 1 0.014″ OK 2 0.012″ OK 3 0.036″ OK B 10.021″ OK 2 0.018″ OK (See FIG. 8) 3 0.030″ OK

As shown in FIGS. 5 through 8, the metallurgical evaluation of thecut-up mounts revealed microstructures with acceptable fusion andpenetration of the Inconel 625 weld material 51 with the base metal 53.No defects such as cracks, major porosity, or lack of fusion were foundin the weld, the interface, and the base metal. The heat-affected zonewas examined carefully, and no micro-cracks were observed.

While the invention has been described with reference to a preferredembodiment, it will be understood by those skilled in the art thatvarious changes may be made and equivalents may be substituted forelements thereof without departing from the scope of the invention. Inaddition, many modifications may be made to adapt to a particularsituation or material to the teachings of the invention withoutdeparting from the essential scope thereof. Therefore, it is intendedthat the invention not be limited to the particular embodiment disclosedas the best mode contemplated for carrying out this invention, but thatthe invention will include all embodiments falling within the scope ofthe appended claims.

1. A method for resurfacing a leading edge of an airfoil comprising thesteps of: removing material from the leading edge of an airfoil;preparing at least the airfoil leading edge for welding; selecting aweld path using an image system; determining welding parameters in orderto avoid cracking; and laser cladding filler material at least onto theairfoil leading edge, and along a selected weld path, using powderfusion welding of superalloy powder and the determined weldingparameters.
 2. The method according to claim 1 wherein the step of lasercladding filler material further comprises laser cladding fillermaterial in a high stress region of the airfoil.
 3. The method accordingto claim 1 wherein the substrate comprises Inconel
 713. 4. The methodaccording to claim 1 wherein the filler material comprises Inconel 625.5. The method according to claim 1 wherein the laser cladding compriseslaser cladding with a CO₂ laser.
 6. The method according to claim 1wherein the laser cladding step further comprises laser cladding with anenergy between about 50 and about 2500 watts.
 7. The method according toclaim 1 wherein the laser cladding step further comprises laser claddingwherein the laser directs a beam on the workpiece with an area ofbetween about 0.02 and about 0.100 square inches.
 8. The methodaccording to claim 1 wherein the laser cladding step further compriseslaser cladding wherein the laser traverses the workpiece surface at arate of between about 3 and about 22 inches per minute.
 9. The methodaccording to claim 1 wherein the laser cladding step deposits fillermaterial in a series of deposition steps.
 10. The method according toclaim 1 wherein the laser cladding step follows a single pass.
 11. Themethod according to claim 1 wherein the laser cladding step follows astitching pattern.
 12. A method for resurfacing a damaged leading edgeof a turbine blisk airfoil comprising the steps of: machining materialaway from the damaged leading edge to a selected height and depthwhereby a machined area is formed; inspecting the machined area byfluorescent penetrating inspection; preparing at least the machined areafor welding; determining a weld path with a laser image system on themachined area; performing a laser fusion welding of at least themachined area with a superalloy powder filler and a CO₂ laser; andautomatically controlling material deposition, energy, and laser travelvelocities during the laser fusion welding to minimize heat cracking inthe airfoil.
 13. The method according to claim 12 wherein the step ofperforming a laser fusion welding further comprises performing a laserfusion welding across a high stress region of an airfoil.
 14. The methodaccording to claim 12 further comprising measuring the depth of thedeposition and repeating a laser fusion welding until a desiredthickness is achieved.
 15. The method according to claim 12 wherein thesuperalloy powder filler comprises Inconel 625 superalloy powder. 16.The method according to claim 12 wherein the step of machining materialfurther comprises machining material to a selected height and depth soas to remove damaged portions of the leading edge.
 17. The methodaccording to claim 12 wherein the step of inspecting the airfoil furthercomprises inspecting in order to confirm the absence of cracks thatwould disqualify the airfoil from repair.
 18. The method according toclaim 12 further comprising the step of machining a repaired airfoil toa desired contour.
 19. The method according to claim 18 furthercomprising a rough machining.
 20. The method according to claim 18further comprising a final machining by hand blending.
 21. The methodaccording to claim 12 wherein the step of performing a laser fusionwelding further comprises laser fusion welding with a co-axial powderfeeder.
 22. A resurfaced airfoil comprising: an airfoil integrallyconnected to a blisk, the airfoil including at least a leading edge, atrailing edge, a top edge, a substrate region, and a repaired region;wherein the repaired region: is welded by powder fusion repair to thesubstrate region, and extends from a welding surface to the airfoilleading edge and from the welding surface to the airfoil top edge. 23.The resurfaced airfoil according to claim 22 wherein the repaired regionis formed by overlapping laser cladding depositions of powdered alloy.24. The resurfaced airfoil according to claim 22 wherein the repairedregion crosses a high stress region of the airfoil.
 25. The resurfacedairfoil according to claim 22 wherein said repaired region furthercomprises a superalloy.
 26. The resurfaced airfoil according to claim 25wherein said repaired region further comprises Inconel
 625. 27. Theresurfaced airfoil according to claim 22 wherein said substrate regionfurther comprises a superalloy.
 28. The resurfaced airfoil according toclaim 27 wherein said substrate region further comprises Inconel 713.29. The resurfaced airfoil according to claim 22 wherein the substrateregion further comprises a weld surface at which the repaired region iswelded to the substrate region.
 30. The resurfaced airfoil according toclaim 29 wherein the weld surface is arcuate in shape.
 31. Theresurfaced airfoil according to claim 22 wherein the repaired regionextends beyond the leading edge.
 32. The resurfaced airfoil according toclaim 22 wherein said repaired region and said substrate region furthercomprise a turbine blisk airfoil.